1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a stator vane in an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
In the turbine section, cooling air used for the inter-stage housing is supplied through the stator vane. FIGS. 5 and 6 show two of the prior art methods of providing cooling air through the vane to the inter-stage housing located inside of the inner endwall or shroud of the stator vanes. In FIG. 5, cooling air from the first leg of a serpentine flow cooling circuit formed within the vane is bled off at an end of the first leg and passed through a hole formed within a cover plate. The hole in the cover plate can be sized to regulate the amount and pressure of the cooling air bled of from the first leg. In the FIG. 5 design, the root turn is located within the hot flow path through the vanes. In FIG. 6, the bleed off hole is located within a root turn of the serpentine circuit. In both of the FIG. 5 and FIG. 6 purge air circuits, the cooling air for the front rim cavity is used for the cooling of the airfoil leading edge region. Because of this, the purge air for the inter-stage housing is over-heated and root turn passage increases turbulence within the cooling air and thus induces uncertainties for the root turn losses. Because of the over-heated purge air for the inter-stage housing, the turbine section overheats resulting in a shorter engine life.